Passive transpirational flow acoustically lined guide vane

ABSTRACT

A passive transpirational flow acoustic liner assembly for a gas turbine engine includes a guide vane assembly and a conduit configured to deliver airflow received from the guide vane. The guide vane assembly includes an airfoil having a transpirational flow acoustic liner. The acoustic liner includes a face sheet defining a portion of an outer surface of the airfoil and having a plurality of first apertures, a segmented member coupled to the face sheet and having a plurality of chambers in fluid communication with the outer surface via the plurality of first apertures, a backing sheet having a plurality of apertures and being coupled to the segmented member such that the segmented member is positioned between the face sheet and the backing sheet, and a plenum coupled to the backing sheet opposite the segmented member and fluidly connected to the conduit.

CROSS-REFERENCE TO RELATED APPLICATION(S)

This application claims the benefit of U.S. Provisional Application No.63/065,923, filed Aug. 14, 2020 for “ACTIVE FLOW CONTROL TRANSPIRATIONALFLOW ACOUSTICALLY LINED GUIDE VANE” by N.D. Sawyers-Abbott and D.Prasad.

BACKGROUND

The present disclosure is related generally to noise attenuation ingeared turbofan engines and more specifically to noise reductionfeatures provided on guide vanes.

Aft fan noise is the dominant source of noise in geared turbofan enginesand acoustic liners provided in the nacelle and engine are the primarymeans for reducing aft fan noise. Acoustic liners provided in the engineitself are becoming more important as nacelles become shorter relativeto fan diameter with the development of turbofans having increasedengine bypass ratios. The addition of acoustic liners to fan exit guidevanes in the fan case can increase the acoustically-treated area in theengine. This can be particularly advantageous in systems that extractair from the bypass duct for cooling core components and for activeclearance control systems, as outlet holes used to extract bypass aircause acoustically-treated area losses. Additional improvements areneeded to increase acoustically-treated area and reduceacoustically-treated area losses caused by cooling systems to enablefurther reductions in nacelle length.

SUMMARY

A passive transpirational flow acoustic liner assembly for a gas turbineengine includes a guide vane assembly and a conduit configured todeliver airflow received from the guide vane. The guide vane assemblyincludes an airfoil having a transpirational flow acoustic liner. Theacoustic liner includes a face sheet defining a portion of an outersurface of the airfoil and having a plurality of first apertures, asegmented member coupled to the face sheet and having a plurality ofchambers in fluid communication with the outer surface via the pluralityof first apertures, a backing sheet having a plurality of apertures andbeing coupled to the segmented member such that the segmented member ispositioned between the face sheet and the backing sheet, and a plenumcoupled to the backing sheet opposite the segmented member and fluidlyconnected to the conduit.

A method for providing acoustic attenuation in a fan section of a gasturbine engine drawing airflow through an acoustic liner on a guidevane, with the acoustic liner being open to an outer flow surface of theguide vane, and exhausting airflow from the guide vane to an area havinga lower pressure than the outer flow surface.

A vane for use in a gas turbine engine includes an airfoil having asuction side and a pressure side and a transpirational flow acousticliner disposed in the airfoil. The liner includes a face sheet defininga portion of an outer surface of the airfoil and having a plurality offirst apertures, a segmented member coupled to the face sheet andincluding plurality of chambers in fluid communication with the outersurface via the plurality of first apertures, and a backing sheet havinga plurality of second apertures coupled to the segmented member suchthat the segmented member is positioned between the face sheet and thebacking sheet, and a plenum coupled to the backing sheet opposite thesegmented member, the plenum in fluid communication with the outersurface via the plurality of second apertures, the plurality of chamber,and the plurality of first apertures.

The present summary is provided only by way of example, and notlimitation. Other aspects of the present disclosure will be appreciatedin view of the entirety of the present disclosure, including the entiretext, claims and accompanying figures.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a quarter-sectional view of a gas turbine engine.

FIG. 2 is a simplified side sectional view of a passive transpirationalacoustic liner assembly and active clearance control system in a gasturbine engine.

FIG. 3 is a cross-sectional view of a fan exit guide vane with atranspirational flow acoustic liner taken along the 3-3 line of FIG. 2.

FIG. 4 is a simplified side sectional view of an alternative passivetranspirational acoustic liner assembly in a gas turbine engine.

While the above-identified figures set forth one or more embodiments ofthe present disclosure, other embodiments are also contemplated, asnoted in the discussion. In all cases, this disclosure presents theinvention by way of representation and not limitation. It should beunderstood that numerous other modifications and embodiments can bedevised by those skilled in the art, which fall within the scope andspirit of the principles of the invention. The figures may not be drawnto scale, and applications and embodiments of the present invention mayinclude features and components not specifically shown in the drawings.

DETAILED DESCRIPTION

A passive flow control transpirational flow acoustically lined guidevane can improve aft fan noise attenuation. In the disclosedtranspirational flow acoustic liner, airflow can be drawn into the vaneby a pressure differential. Suction created on the acoustic liner flowsurface can reduce drag on the guide vane by retaining laminar flowacross the rougher acoustic liner surface of the vane. Airflow drawnthrough the vane can be exhausted to an outer surface of a nacellehaving a lower pressure than the fan bypass duct, or can be used forcooling core components and for active clearance control in the turbine.

FIG. 1 is a quarter-sectional view of a gas turbine engine 20 thatincludes fan section 22, compressor section 24, combustor section 26 andturbine section 28. Fan section 22 includes fan 42, exit guide vane 64,and fan case 65. Fan section 22 drives air along bypass flow path Bwhile compressor section 24 draws air in along core flow path C whereair is compressed and communicated to combustor section 26. In combustorsection 26, air is mixed with fuel and ignited to generate a highpressure exhaust gas stream that expands through turbine section 28where energy is extracted and utilized to drive fan section 22 andcompressor section 24.

Although the disclosed non-limiting embodiment depicts a turbofan gasturbine engine, it should be understood that the concepts describedherein are not limited to use with turbofans as the teachings may beapplied to other types of turbine engines; for example a low-bypassturbine engine, or a turbine engine including a three-spool architecturein which three spools concentrically rotate about a common axis andwhere a low spool enables a low pressure turbine to drive a fan via agearbox, an intermediate spool that enables an intermediate pressureturbine to drive a first compressor of the compressor section, and ahigh spool that enables a high pressure turbine to drive a high pressurecompressor of the compressor section.

The example engine 20 generally includes low speed spool 30 and highspeed spool 32 mounted for rotation about an engine central longitudinalaxis A relative to an engine static structure 36 via several bearingsystems 38. It should be understood that various bearing systems 38 atvarious locations may alternatively or additionally be provided.

Low speed spool 30 generally includes inner shaft 40 that connects fan42 and low pressure compressor section 44 to low pressure turbinesection 46. Inner shaft 40 drives fan 42 through a speed change device,such as geared architecture 48, to drive fan 42 at a lower speed thanlow speed spool 30. High-speed spool 32 includes outer shaft 50 thatinterconnects high pressure compressor section 52 and high pressureturbine section 54. Inner shaft 40 and outer shaft 50 are concentric androtate via bearing systems 38 about engine central longitudinal axis A.

Combustor 56 is arranged between high pressure compressor 52 and highpressure turbine 54. In one example, high pressure turbine 54 includesat least two stages to provide a double stage high pressure turbine 54.In another example, high pressure turbine 54 includes only a singlestage. As used herein, a “high pressure” compressor or turbineexperiences a higher pressure than a corresponding “low pressure”compressor or turbine.

The example low pressure turbine 46 has a pressure ratio that is greaterthan about 5. The pressure ratio of the example low pressure turbine 46is measured prior to an inlet of low pressure turbine 46 as related tothe pressure measured at the outlet of low pressure turbine 46 prior toan exhaust nozzle.

Mid-turbine frame 58 of engine static structure 36 is arranged generallybetween high pressure turbine 54 and low pressure turbine 46.Mid-turbine frame 58 further supports bearing systems 38 in turbinesection 28 as well as setting airflow entering low pressure turbine 46.

The core airflow C is compressed by low pressure compressor 44 then byhigh pressure compressor 52, mixed with fuel, and ignited in combustor56 to produce high speed exhaust gases that are then expanded throughhigh pressure turbine 54 and low pressure turbine 46. Mid-turbine frame57 includes airfoils/vanes 60, which are in the core airflow path andfunction as an inlet guide vane for low pressure turbine 46. Utilizingvanes 60 of mid-turbine frame 58 as inlet guide vanes for low pressureturbine 46 decreases the length of low pressure turbine 46 withoutincreasing the axial length of mid-turbine frame 58. Reducing oreliminating the number of vanes in low pressure turbine 46 shortens theaxial length of turbine section 28. Thus, the compactness of gas turbineengine 20 is increased and a higher power density may be achieved.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than ten (10), the gearedarchitecture 48 is an epicyclic gear train, such as a planetary gearsystem or other gear system, with a gear reduction ratio of greater thanabout 2.3 and the low pressure turbine 46 has a pressure ratio that isgreater than about 5. In one disclosed embodiment, the engine 20 bypassratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout 5:1. Low pressure turbine 46 pressure ratio is pressure measuredprior to inlet of low pressure turbine 46 as related to the pressure atthe outlet of the low pressure turbine 46 prior to an exhaust nozzle.The geared architecture 48 may be an epicycle gear train, such as aplanetary gear system or other gear system, with a gear reduction ratioof greater than about 2.5:1. It should be understood, however, that theabove parameters are only exemplary of one embodiment of a gearedarchitecture engine and that the present invention is applicable toother gas turbine engines including direct drive turbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft, withthe engine at its best fuel consumption—also known as “bucket cruiseThrust Specific Fuel Consumption (‘TSFC’)”—is the industry standardparameter of lbm of fuel being burned divided by lbf of thrust theengine produces at that minimum point. “Low fan pressure ratio” is thepressure ratio across the fan blade alone, without a fan exit guide vane64 system. The low fan pressure ratio as disclosed herein according toone non-limiting embodiment is less than about 1.45. “Low corrected fantip speed” is the actual fan tip speed in ft/sec divided by an industrystandard temperature correction of [(Tambient deg R)/518.7){circumflexover ( )}0.5]. The “Low corrected fan tip speed” as disclosed hereinaccording to one non-limiting embodiment is less than about 1150ft/second.

FIG. 2 illustrates passive transpirational flow acoustic liner assembly61 in a fan section of the gas turbine engine of FIG. 1. FIG. 2 showsfan section 22 with fan blade 42, fan rotor 62, fan exit guide vane 64,and fan case 65; nacelle 66 having inlet cowl 67 with outer surface 68,fan cowl door 70, and thrust reverser 72; core housing 74, inner fixedstructure (IFS) 75, cavity 76, guide vane manifold 78, conduit 80, andactive clearance control (ACC) system 82 including valve 84; andcontroller 86. Nacelle 66, including inlet cowl 67, fan cowl door 70,thrust reverser 72, and IFC 75 forms a shroud around the engine. Theseelements are present but not shown in FIG. 1. Fan case 65 is spacedradially outwardly of fan blade 42. Fan cowl door 70 forms a portion ofnacelle 66 positioned around fan case 65. Thrust reverser 72 forms aportion of nacelle 66 aft of fan cowl door 70. Exit guide vane 64 ispositioned aft of fan blade 42 and is secured to fan case 65 at aradially outer end and core housing 74 at a radially inner end.

Exit guide vane 64 includes an airfoil having base 88 (radiallyinnermost end), tip 90 (radially outermost end), leading edge 92,trailing edge 94, pressure side 96 (shown in FIG. 3), suction side 98,acoustic liner 100, and channel 101. Exit guide vane 64 extends betweencore housing 74 and fan case 65 with base 88 fixed to core housing 84and tip 90 fixed to fan case 65. Exit guide vane 64 is one of multiplecircumferentially spaced guide vanes. Exit guide vanes 64 remove theswirl imparted to the bypass flow by fan blades 42 and straighten orredirect flow in a substantially axial direction. As illustrated in FIG.2, leading edge 92 can be swept rearwardly over a full radial extent ofthe bypass duct from base 88 to tip 90 to reduce noise. The arrangementand number of guide vanes 64 can be optimized to improve noiseattenuation as taught in U.S. Pat. No. 1,010,719.

Acoustic liner 100 is a transpirational flow liner open to an outersurface of guide vane 64 and to channel 101 thereby fluidly connectingchannel 101 with bypass flow B. Channel 101 is connected to acousticliner 100 at base 88 of exit guide vane 64. Channel 101 is fluidlyconnected to conduit 80 in core housing 74 via guide vane manifold 78.Conduit 80 extends through cavity 76 with one or more outlets located incavity 76 as described further below. As such, cavity 76 is fluidlyconnected to the fan bypass duct and bypass flow B via conduit 80, guidevane manifold 78, channel 101, and acoustic liner 100. Acoustic liner100 can extend a substantially full length of guide vane 64 from base 88to tip 90 to provide maximum acoustic benefit or can extend a maximumlength that can be accommodated by guide vane 64 without compromisingstructural integrity.

A pressure differential between the fan bypass duct at an outer surfaceguide vane 64 and cavity 76 causes bypass flow B to be drawn into exitguide vane 64 through acoustic liner 100 and conduit 80 and into cavity76. The biased flow through exit guide vane 64 creates suction on theouter flow surface of exit guide vane 64. The suction on the outer flowsurface helps prevent separation of flow and retain a laminar flow overthe acoustic liner surface, which has increased surface roughness incomparison to the remainder of the outer surface of exit guide vane 64.Biased flow through acoustic liner 100 thereby reduces drag on exitguide vane 64, which can lead to improved TSFC. Additional acousticbenefit can be gained with biased flow through acoustic liner 100 asflow through acoustic liner 100 can provide enhanced acousticdissipation. Sound waves drawn into acoustic liner 100 with airflow exitwith airflow from the fan duct thereby reducing a potential for noise toreenter the fan duct.

Guide vane manifold 78 can receive cooling air from the plurality ofguide vanes 64 via channels 101. Guide vane manifold 78 can be anannular or partial ring (shown in phantom) connecting outlets ofchannels 101 from the plurality of guide vanes 64 having acoustic liners100. Guide vane manifold 78 can be any shape and configuration capableof combining cooling air flow received from guide vanes 64. Conduit 80connects to guide vane manifold 78.

Conduit 80 can connect guide vane manifold 78 to cavity 76 via one ormore outlets or passages 102 and/or ACC system 82. Valve 84 can controlflow through conduit 80.

Cooling air can be delivered from conduit 80 to core cavity 76 forcooling electrical and mechanical components, replacing the air incavity 76, and feeding ACC system 82. Flammability regulations requirethat air in cavity 76 be replaced to remove flammable gases. Passivetranspirational flow acoustic liner assembly 61 can be configured toreplace core cavity air greater than one time every minute with airreceived from guide vanes 64. One or more tubes or passages 102 canconnect to conduit 80 and direct cooling air to specified locations incore cavity 76. For example, cooling air can be directed to electricalpower feeder cables (PFC) or an aircraft's environmental control system(ECS) valve housed in cavity 76 and also to an integrated drivegenerator (IDG), among other components that require cooling duringvarious modes of operation. Use of passages 102 to cool components canreduce or eliminate the need for providing holes through inner fixedstructure (IFS) 75, which are typically provided to cool components andreplace the air in cavity 76. Holes through IFS 75 reduce theacoustically-treated area of the fan bypass duct. As such,acoustically-treated area can be increased by delivering cooling airreceived from guide vanes 64 to cavity 76. In addition, the penetrationsthrough IFS 75 also represent additional loss mechanisms in the TSFC dueto the shape of the penetrations through the IFS. One or more passages102 (shown in phantom) can extend circumferentially around the enginecore and can include a plurality of holes directed at components or canbranch off from conduit 80 or a duct from conduit 80 at one or morelocations and extend to identified hot regions or components requiringcooling. Cooling air can be circulated in cavity 76 and exhausted fromcavity 76 to atmosphere through aft vent 105. A pressure differentialbetween cavity 76 and the external atmosphere drives cooling air outthrough aft vent 105.

ACC system 82 can be used to control a clearance between turbine rotorblade tips and a shroud or blade outer air seals (BOAS, not shown) asknown in the art. ACC system 82 includes conduit 80, valve 84, and oneor more annular turbine cooling passages 104. Valve 82 is configured tocontrol flow through conduit 80. Cooling air from guide vanes 64 can bedelivered to cooling passages 104 via conduit 80. ACC system 82 can beconfigured as known in the art to deliver cooling air to one or morecooling passages 104 surrounding turbine rotor blades to causecontraction a thermal control ring (not shown) toward blade tips toreduce losses at blade tips and improve engine efficiency. Coolingpassages 104 can be mounted to the high pressure turbine section 54where operating temperatures are highest or can be mounted to both highpressure turbine section 54 and low pressure turbine section 46. Coolingpassages 104 can direct cooling air received from guide vanes 64 througha plurality of holes in cooling passages 104 disposed circumferentially(shown in phantom) around the turbine section to a back side (radiallyouter surface) of a thermal control ring surrounding the rotor blades.The cooling air can cause thermal contraction of the thermal controlring inward toward rotor blade tips thereby reducing a clearance betweenthe blade tips the thermal control ring or BOAS connected thereto asknown in the art. Cooling passages 104 can be connected by a manifoldand/or accumulator as known in the art to divide flow among differentturbine stages. One or more valves (not shown) can be used to furtherregulate flow. For example, a valve or accumulator can be locatedbetween turbine stages. Cooling air feeding ACC system 84 and impingingon the thermal control ring is fluidly connected to cavity 76 and can beexhausted through aft vent 105.

Valve 84 can be any type of valve known in the art capable of providingon/off flow control or regulated flow control through conduit 80. Valve84 can be controlled by controller 86 (e.g., engine FADEC system). Valve84 is used to operate ACC system 82. An open/closed position of valve 84can be scheduled based on power ratings or can be regulated based onneed, for example, as determined by sensors as known in the art. Valve84 can be closed, for example, during aircraft takeoff climb duringwhich time ACC system 82 is generally not necessary. High operatingtemperatures during aircraft takeoff and climb cause thermal growth ofturbine blades toward the shroud or BOAS thereby minimizing blade tipclearance and reducing efficiency losses associated with tip leakage.There is an increased need to cool components in cavity 76 duringtakeoff and climb. With valve 84 in a closed position, all cooling airextracted from the fan bypass duct through guide vane 64 can be directedto cavity 76 via one or more passages 102.

Valve 84 can be opened during cruise as power settings are reduced,operating temperatures drop, and blade tip clearance increases. Duringcruise, cooling air can be directed to ACC system 82 to cause thermalcontrol ring to contract toward turbine blade tip thereby minimizing tipclearance and reducing efficiency losses associated with tip leakage. Aportion of cooling air through conduit 80 can be directed to cavity 76through one or more passages 102 to cool components and replace air incavity 76. The need for cooling components is reduced during cruise asoperating temperatures are lower than during takeoff and climbingoperations. Conduit 80 and passages 102 and 104 can be sized andconfigured to provide sufficient cooling air flow to ACC system 82 whenvalve 84 is open. In some embodiments, cooling holes through IFS 75 canbe used to provide additional cooling to components in cavity 76.

Typically, bypass air for ACC systems is provided through a largeopening in core housing 74. Such opening reduces acoustically-treatedarea. The use of passive transpirational acoustic flow assembly 61 forACC system 82 eliminates the need to provide an opening in core housing74 thereby reducing acoustically-treated area losses. Additionally, inlow fan pressure ratio systems, a hooded scoop on core housing 74 can berequired to direct cooling air into the core housing opening for the ACCsystem. Such hooded scoop extends into the fan bypass duct where it canincrease drag and thereby reduce engine efficiency. The need for ahooded scoop is also eliminated when passive transpirational acousticflow assembly 61 is used to provide cooling air for ACC system 82.

Guide vane manifold 78, conduit 80, cavity passages 102, valve 84, andACC system passages 104 can be formed from materials suitable for atemperature and environmental conditions in which they are located.Suitable materials can include but are not limited to compositematerials, plastics, aluminum, steel, and titanium. Conduit 80 can bedivided into sections formed of differing materials based on variationsin operational temperatures between the location of guide vane 64 andturbine section 28.

FIG. 3 illustrates a cross-sectional view of fan exit guide vane 64 withacoustic liner 100 taken along the 3-3 line of FIG. 2. FIG. 3 showsleading edge 92, trailing edge 94, pressure side surface 98, suctionside surface 96, and acoustic liner 100. Acoustic liner 100 includesperforated face sheet 106, segmented member 108 with chambers 110,perforated backing sheet 112, and plenum 114. Acoustic liner 100 can bea cartridge-type member that can be inserted into a pocket of guide vane64 or can be integrally formed with guide vane 64. Acoustic liner 100can be formed via an additive manufacturing process such as powder bedmetallurgy, direct metal laser sintering, select laser sintering, selectlaser melting, electron beam melting or other. Acoustic liner 100 can bedisposed on suction side 98 such that perforated face sheet 106 forms aportion of the suction side outer surface. Most flow separation andlosses occur on the low pressure or suction side of the airfoil. Assuch, greater benefit may be achieved by improving flow efficiency onsuction side 98. In alternative embodiments, acoustic liner 100 can beprovided on both suction side 98 and pressure side 96 or pressure side96, alone.

Perforated face sheet 106 can define a portion of the outer surface ofguide vane 64. Segmented member 108 can be coupled to a back side ofperforated face sheet 106. Perforated backing sheet 112 can be coupledto segmented member 108 such that segmented member 108 is positionedbetween perforated face sheet 106 and perforated backing sheet 112.Plenum 114 can be coupled to perforated backing sheet 112 oppositesegmented member 108. As arranged, plenum 114 is in fluid communicationwith the outer surface of guide vane 64 via openings in perforatedbacking sheet 112, chambers 110 in segmented member 108, and openings inperforated face sheet 106, such that bypass flow B can enter exit guidevane 64 through perforated face sheet 106 and can flow through segmentedmember 108 and perforated backing sheet 112 to plenum 114. Plenum 114 isconnected to or forms channel 101 open to base 88, which is fluidlyconnected guide vane manifold 78 and conduit 80. Plenum 114 can beclosed at tip 90.

Perforated face sheet 106 includes a plurality of apertures 116configured to allow bypass flow B into guide vane 64 and communicateacoustic energy to underlying chambers 110 of segmented member 108.Apertures 116 can be round holes having a diameter in the range of 3 to50 thousandths of an inch and can cover at least 5% of the surface offace sheet 106. Apertures 116 can be configured to create suction alongthe outer surface of guide vane 64 to prevent boundary layer flowseparation, promote laminar flow, and reduce skin friction losses acrossexit guide vane 64 in the bypass duct aft of fan 42. Apertures 116 canbe arranged in rows or patterns with substantially uniform spacingbetween apertures 116. The number, size, and arrangement of apertures116 can be designed to maximize acoustic attenuation performance whileoptimizing suction along the outer surface of exit guide vane 64 toenhance flow and minimize drag. Apertures 116 can be sized to limitclogging. In alternative embodiments, apertures 116 can have othershapes and configurations designed to optimize flow along the outersurface of exit guide vane 64 and/or optimize acoustic attenuation.Perforated face sheet 106 can be configured to provide a substantiallysmooth surface with suction side 98 of exit guide vane 64. Perforatedface sheet 106 can be constructed of metal, composites, and/or otherknown materials.

Segmented member 108 is a cellular structure such as a honeycombstructure as known in the art with open chambers 110. Segmented member108 is disposed on the back side of perforated face sheet 106 such thatperforated face sheet 106 covers chambers 110 and apertures 116 are opento chambers 110. Segmented member 108 can be designed to providestructural support. The size of chambers 110 can be optimized forproviding structural support and for receiving acoustic energy.Segmented member 108 can be constructed of metal, composites, and/orother known materials. In some embodiments, acoustic liner 100 can be adouble degree of freedom (DDOF) liner (not shown) in which segmentedmember 108 includes two cellular structure layers separated by aperforated sheet allowing air flow between the cellular structures. DDOFliners can provide additional acoustic attenuation benefit over a singlerow of honeycomb resonators. The first and second cellular structurelayers can be tuned to different tones to achieve improved acousticattenuation.

Perforated backing sheet 112 is disposed on a back side of segmentedmember 108 such that segmented member 108 is sandwiched betweenperforated face sheet 106 and perforated backing sheet 112. Perforatedbacking sheet 112 provides structural support for segmented member 108and allows bypass flow B to be drawn through acoustic liner 100.Perforated backing sheet 112 is arranged to cover chambers 110 on theback side of segmented member. Perforated backing sheet 112 includes aplurality of apertures 118 configured to allow bypass flow B into plenum114. Apertures 118 are open to chambers 110 and communicate bypass flowB from chambers 110 in segmented member 108 to plenum 114. Apertures 118can be round holes having a diameter in the range of 3 to 100thousandths of an inch and can cover at least 5% of the surface ofperforated backing sheet 112. Apertures 118 can be configured to provideoptimal flow dynamics through acoustic liner 100. Apertures 118 can bearranged in rows or patterns with substantially uniform spacing betweenapertures 118. The number, size, and arrangement of apertures 118 can bedesigned to optimize fluid flow through acoustic liner 100. Inalternative embodiments, apertures 118 can have other shapes andconfigurations designed to optimize fluid flow. Perforated backing sheet112 can be constructed of metal, composites, and/or other knownmaterials.

Plenum 114 is disposed on a back side of perforated backing sheet 112.Plenum 114 can be formed by an enclosure connected to perforated backingsheet 112 or can be a gap formed between perforated backing sheet 112and an interior surface of guide vane 64 upon assembly. Plenum 114 canbe closed at a radially outer end of acoustic liner 100 near tip 90 andcan be open at a radially inner end of acoustic liner 100 near base 88where plenum 114 forms or connects to channel 101 to allow bypass flow Bto be drawn through acoustic liner 100 into guide vane manifold 78 andconduit 80. Plenum 114 can be sized to provide effective fluid flowthrough acoustic liner 100.

An acoustic liner 100 can be included in every fan exit guide vane 64 tomaximize the size of the acoustically-treated area within the engine. Inalternative embodiments, acoustic liners 100 can be incorporated in lessthan all of exit guide vanes 64. For example, an acoustic liner 100 canbe provided in every other exit guide vane 64 such that approximatelyhalf of the exit guide vanes 64 have acoustic liners 100 andapproximately half of exit guide vanes 64 do not have acoustic liners100. The number and arrangement of exit guide vanes 64 with acousticliners 100 can be designed to optimize noise attenuation.

In some embodiments, acoustic attenuation properties of acoustic liner100 can be optimized as described in U.S. Pat. No. 7,540,354,“Micro-perforated Acoustic Liner,” and U.S. Pat. No. 10,107,191, “GearedGas Turbine Engine with Reduced Fan Noise,” which are incorporated byreference.

During operation of transpirational flow acoustic liner assembly 61,bypass flow B is drawn into exit guide vane 64 aft of fan blades 42,thereby establishing suction on an outer surface of face sheet 106 ofacoustic liner 100. Suction is created by drawing bypass flow B throughacoustic liner 100 into plenum 114, guide vane manifold 78, conduit 80,and cavity 76. A pressure differential between the fan bypass duct (atguide vane 64) and core cavity 76 causes bypass flow B to be drawn intoguide vane 64 and through conduit 80 to cavity 76 via passage 102 or ACCsystem 82. Controller 86 can be used to preferentially regulate flowthrough acoustic liner 100 during operation of the gas turbine engine byopening, closing, or modulating a position of valve 84 to regulate flow.For example, valve 84 can be closed at takeoff, causing cooling air toenter cavity 76 through one or more passages 102 to cool components. Atcruise, valve 84 can be opened, causing cooling air to be directed toACC system 82 to improve engine efficiency by reducing tip leakage.During ACC operation, a portion of the cooling air can continue to bedelivered to cavity 76 via one or more passages 102 to cool components.

FIG. 4 illustrates passive transpirational flow acoustic liner assembly161 in a fan section of a gas turbine engine. FIG. 4 shows portions ofthe gas turbine engine as described with respect to FIG. 2, includingfan section 22, fan blade 42, fan rotor 62, fan case 65, nacelle 66,inlet cowl 67, outer surface 68, fan cowl door 70, thrust reverser 72,core housing 74, and IFC 75. Acoustic liner assembly 161 includes fanexit guide vane 164, guide vane manifold 178, conduit 180, and exhaustoutlet 182. Conduit 180 is provided in nacelle 66. Conduit 180 canextend through fan case 65 into the cavity between fan case 65 and thenacelle outer surface. Conduit 180 can connect to exit guide vane 164 atone end via manifold 178 and exhaust outlet 182 at an opposite end.Exhaust outlet 182 can be located on outer surface 68 of nacelle 66 and,specifically, at an area of lower pressure than fan duct pressure B.Exit guide vane 164 is positioned aft of fan blade 42 and is secured tofan case 65.

Exit guide vane 164 includes an airfoil having base 188 (radiallyinnermost end), tip 190 (radially outermost end), leading edge 192,trailing edge 194, pressure side 196 (not shown), suction side 198,acoustic liner 200, and channel 201. Exit guide vane 164 extends betweencore housing 84 and fan case 65 with base 188 fixed to core housing 84and tip 190 fixed to fan case 65. Exit guide vane 164 is one of multiplecircumferentially spaced guide vanes. Exit guide vanes 164 remove theswirl imparted to the bypass flow by fan blades 42 and straighten orredirect flow in a substantially axial direction. As described withrespect to FIG. 2 and illustrated in FIG. 4, leading edge 192 can beswept rearwardly over a full radial extent of the bypass duct from base188 to tip 190 to reduce noise and the arrangement and number of guidevanes 164 can be optimized to improve noise attenuation.

Acoustic liner 200 is a transpirational flow acoustic liner and can besubstantially the same as acoustic liner 100 of FIG. 3 with theexception that channel 201 extends from plenum 114 tip 190 as opposed tobase 188. Acoustic liner 200 (plenum 114) can be closed at base 188.Acoustic liner 200 is open to an outer surface of guide vane 64 and tochannel 201 thereby fluidly connecting channel 201 with bypass flow B.Channel 201 is connected to acoustic liner 200 at tip 190 of exit guidevane 64. Channel 201 is fluidly connected to conduit 180 of fan case 65via guide vane manifold 178, such that conduit 180 is fluidly connectedto the bypass duct and bypass flow B. Acoustic liner 200 can extend asubstantially full length of guide vane 164 from base 188 to tip 190 toprovide maximum acoustic benefit, or can extend a maximum length thatcan be accommodated by guide vane 164 without compromising structuralintegrity.

A pressure differential between the bypass duct at an outer surface ofguide vane 164 and outlet 182 causes bypass flow B to be drawn into exitguide vane 164 through acoustic liner 200 and can exhaust throughconduit 180 and outlet 182 to outer surface 68 of nacelle 66. The biasedflow through exit guide vane 164 creates suction on the outer flowsurface of exit guide vane 164. The suction on the outer flow surfacehelps prevent separation of flow and retain a laminar flow over theacoustic liner surface, which has increased surface roughness incomparison to the remainder of the outer surface of exit guide vane 164.Biased flow through acoustic liner 200 thereby reduces drag on exitguide vane 164, which can lead to improved TSFC. Additional acousticbenefit can be gained with biased flow through acoustic liner 200 asflow through acoustic liner 200 can provide enhanced acousticdissipation.

Guide vane manifold 178 can receive cooling air from the plurality ofguide vanes 164 via channels 201. Guide vane manifold 178 can be anannular or partial ring connecting outlets of channels 201 from theplurality of guide vanes 164 having acoustic liners 200. Guide vanemanifold 178 can be any shape and configuration capable of combiningcooling air flow received from guide vanes 164. Conduit 180 connects toguide vane manifold 178.

Air flow through acoustic liner 200 can be channeled through passage 180and exhausted through outlet 182 on outer surface 68 of nacelle 66forward of thrust reverser 72 to promote a laminar flow or reduce skinfriction losses along outer surface 68. Exhaust outlet 182 can bepositioned at any area of lower pressure than fan duct pressure B.- Insome embodiments, exhaust outlet 182 can include a plurality of openingsdisposed about the circumference nacelle 66. In some embodiments,exhaust of air flow can be limited to outer surface 68 on an upper ortop side of nacelle 66 where increased flow efficiency can provideincreased benefit. Depending on the amount of air flow drawn throughacoustic liner 200 and exiting outlet 182, drag reduction of exhaustflow E may be achieved on outer surface 68. Drag reduction across outersurface 68 can reduce drag and thereby further improve fuel burn.

During operation of transpirational flow acoustic liner assembly 161,bypass flow B is drawn into exit guide vane 164 aft of fan blades 42,thereby establishing suction on an outer surface of face sheet 106 ofacoustic liner 200. Suction is created by drawing bypass flow B throughacoustic liner 200 into plenum 114 and conduit 180. A pressuredifferential between the bypass duct (at guide vane 164) and atmosphere(exhaust outlet 182) causes bypass flow B to be drawn into guide vane164 and through conduit 180 and out through exhaust outlet 182.

The disclosed passive transpirational flow acoustic liner assemblies 61and 161 can increase the acoustically-treated area of the fan bypassduct and reduce drag on exit guide vanes 64 and 164, respectively, bypromoting drag reduction thereby improving fuel burn. Passivetranspirational flow acoustic liner assembly 161 can further promotedrag redution on nacelle outer surface 68. Improved noise attenuationresulting from an increase in acoustically-treated area can allow forreductions in fan duct length, which can enable installation of largerbypass ratio engine systems. Passive transpirational flow acoustic linerassembly 61 can further increase the acoustically-treated area byeliminating or reducing the need for cooling holes through IFS 75 neededto cool components and operate an ACC system, and can reduce drag byeliminating the need for a hooded scoop or other flow guide structurefor operation of an ACC system or cooling holes. Passive transpirationalflow acoustic liner assembly 61 can be used to cool components locatedin core compartment 76 and for ACC in turbine section 28.

Summation

Any relative terms or terms of degree used herein, such as“substantially”, “essentially”, “generally”, “approximately” and thelike, should be interpreted in accordance with and subject to anyapplicable definitions or limits expressly stated herein. In allinstances, any relative terms or terms of degree used herein should beinterpreted to broadly encompass any relevant disclosed embodiments aswell as such ranges or variations as would be understood by a person ofordinary skill in the art in view of the entirety of the presentdisclosure, such as to encompass ordinary manufacturing tolerancevariations, incidental alignment variations, transient alignment orshape variations induced by thermal, rotational or vibrationaloperational conditions, and the like. Moreover, any relative terms orterms of degree used herein should be interpreted to encompass a rangethat expressly includes the designated quality, characteristic,parameter or value, without variation, as if no qualifying relative termor term of degree were utilized in the given disclosure or recitation.

Discussion of Possible Embodiments

The following are non-exclusive descriptions of possible embodiments ofthe present invention.

A passive transpirational flow acoustic liner assembly for a gas turbineengine includes a guide vane assembly and a conduit configured todeliver airflow received from the guide vane. The guide vane assemblyincludes an airfoil having a transpirational flow acoustic liner. Theacoustic liner includes a face sheet defining a portion of an outersurface of the airfoil and having a plurality of first apertures, asegmented member coupled to the face sheet and having a plurality ofchambers in fluid communication with the outer surface via the pluralityof first apertures, a backing sheet having a plurality of apertures andbeing coupled to the segmented member such that the segmented member ispositioned between the face sheet and the backing sheet, and a plenumcoupled to the backing sheet opposite the segmented member and fluidlyconnected to the conduit.

The assembly of the preceding paragraph can optionally include,additionally and/or alternatively, any one or more of the followingfeatures, configurations and/or additional components:

The assembly of the preceding paragraphs, wherein the conduit isdisposed in a core housing of the gas turbine engine and fluidlyconnected to the plenum by an opening at a base of the airfoil.

The assembly of any of the preceding paragraphs, wherein the conduitincludes an exhaust passage located to deliver airflow to componentsdisposed in the core housing.

The assembly of any of the preceding paragraphs, wherein the conduit isconnected to an active clearance control system in a turbine section ofthe gas turbine engine and configured to deliver airflow to the activeclearance control system.

The assembly of any of the preceding paragraphs, wherein the activeclearance control system comprises a valve located in a fluid positionaft of the exhaust passage and configured to regulate the airflow to theturbine section.

The assembly of any of the preceding paragraphs, wherein the conduit isdisposed in a nacelle of the gas turbine engine and fluidly connected tothe plenum by an opening at a tip of the airfoil.

The assembly of any of the preceding paragraphs which further includesan exhaust outlet open to a radially outer surface of the nacelle andfluidly connected to the conduit.

A method for providing acoustic attenuation in a fan section of a gasturbine engine drawing airflow through an acoustic liner on a guidevane, with the acoustic liner being open to an outer flow surface of theguide vane, and exhausting airflow from the guide vane to an area havinga lower pressure than the outer flow surface.

The method of the preceding paragraph can optionally include,additionally and/or alternatively, any one or more of the followingfeatures, configurations, additional components, and/or steps:

The method of the preceding paragraphs, wherein airflow is drawn throughthe acoustic liner and into a plenum of the guide vane.

The method of any of the preceding paragraphs which further includesdrawing the airflow from a base of the guide vane into a conduitdisposed in a core housing of the gas turbine engine.

The method of any of the preceding paragraphs can further includeexhausting the airflow from the conduit to a cavity of the core housingto cool components located in the cavity.

The method of any of the preceding paragraphs which further includesdelivering the airflow to an active clearance control system in aturbine section of the gas turbine engine.

The method of any of the preceding paragraphs which further includesfurther opening and closing a valve in the conduit to regulate coolingairflow to the turbine section for active clearance control, whereinopening the valve allows airflow to the turbine section and closing thevalve inhibits airflow to the turbine section.

The method of any of the preceding paragraphs, wherein the valve isclosed during a takeoff operation of the gas turbine engine and openduring a cruise operation of the gas turbine engine

A vane for use in a gas turbine engine includes an airfoil having asuction side and a pressure side and a transpirational flow acousticliner disposed in the airfoil. The liner includes a face sheet defininga portion of an outer surface of the airfoil and having a plurality offirst apertures, a segmented member coupled to the face sheet andincluding plurality of chambers in fluid communication with the outersurface via the plurality of first apertures, and a backing sheet havinga plurality of second apertures coupled to the segmented member suchthat the segmented member is positioned between the face sheet and thebacking sheet, and a plenum coupled to the backing sheet opposite thesegmented member, the plenum in fluid communication with the outersurface via the plurality of second apertures, the plurality of chamber,and the plurality of first apertures.

The vane of the preceding paragraph can optionally include, additionallyand/or alternatively, any one or more of the following features,configurations and/or additional components:

The vane of the preceding paragraphs, wherein the vane has a tip and abase disposed opposite the tip and wherein the plenum is open at one ofthe tip and the base and closed at the other of the tip and the base.

The vane of any of the preceding paragraphs, wherein first apertures ofthe plurality of first apertures have a diameter in a range of 3 to 50thousandths of an inch.

The vane of any of the preceding paragraphs

The assembly of any of the preceding paragraphs, wherein the secondapertures of the plurality of second apertures have a diameter in therange of 3 to 100 thousandths of an inch

The vane of any of the preceding paragraphs, wherein the plenum is openat the base.

The vane of any of the preceding paragraphs, wherein the liner facesheet is located on the suction side of the airfoil.

While the invention has been described with reference to an exemplaryembodiment(s), it will be understood by those skilled in the art thatvarious changes may be made and equivalents may be substituted forelements thereof without departing from the scope of the invention. Inaddition, many modifications may be made to adapt a particular situationor material to the teachings of the invention without departing from theessential scope thereof. Therefore, it is intended that the inventionnot be limited to the particular embodiment(s) disclosed, but that theinvention will include all embodiments falling within the scope of theappended claims.

The invention claimed is:
 1. A passive transpirational flow acousticliner assembly for a gas turbine engine, the assembly comprising: aguide vane assembly comprising: an airfoil having a suction side and apressure side; and a transpirational flow acoustic liner disposed in theairfoil, the liner comprising: a face sheet defining a portion of anouter surface of the airfoil and having a plurality of first apertures;a segmented member coupled to the face sheet, the segmented membercomprising a plurality of chambers in fluid communication with the outersurface via the plurality of first apertures; a backing sheet coupled tothe segmented member such that the segmented member is positionedbetween the face sheet and the backing sheet, the backing sheet having aplurality of second apertures; and a plenum coupled to the backing sheetopposite the segmented member; and a conduit fluidly connected to theplenum and configured to deliver airflow received from the plenum to oneof: a plurality of exhaust outlets open to a radially outer surface of anacelle of the gas turbine engine and disposed about a partialcircumference of the nacelle; and a plurality of exhaust passagesfluidly connected to the conduit and disposed in a core housing of thegas turbine engine, wherein the passive transpirational flow acousticliner assembly is configured to deliver the airflow without use of apump.
 2. The assembly of claim 1, wherein the conduit is disposed in thecore housing of the gas turbine engine and fluidly connected to theplenum by an opening at a base of the airfoil.
 3. The assembly of claim2, wherein the plurality of exhaust passages extend from the conduit andare located to deliver the airflow to components disposed in the corehousing.
 4. The assembly of claim 3, wherein the conduit is connected toan active clearance control system in a turbine section of the gasturbine engine and configured to deliver the airflow to the activeclearance control system.
 5. The assembly of claim 4, wherein the activeclearance control system comprises a valve located in a fluid positionaft of the plurality of exhaust passages and configured to regulate theairflow to the turbine section, wherein the airflow is delivered to theactive clearance control system when the valve is open and the airflowis delivered to components disposed in the core housing via theplurality of exhaust passages when the valve is closed.
 6. The assemblyof claim 1, wherein the conduit is disposed in the nacelle of the gasturbine engine and fluidly connected to the plenum by an opening at atip of the airfoil.
 7. The assembly of claim 6, wherein the location ofplurality of exhaust outlets is limited to an upper side of the nacelle.8. A method for providing acoustic attenuation in a fan section of a gasturbine engine, the method comprising: drawing airflow through anacoustic liner on a guide vane and into a plenum of the guide vane, theacoustic liner open to an outer flow surface of the guide vane; drawingthe airflow into a conduit fluidly connected to a plenum of the acousticliner; and exhausting the airflow from the guide vane to an area havinga lower pressure than the outer flow surface, wherein exhausting theairflow comprises exhausting the airflow through one of: a plurality ofexhaust outlets open to a radially outer surface of a nacelle of the gasturbine engine and disposed about a partial circumference of thenacelle; and a plurality of exhaust passages fluidly connected to aconduit and disposed in a core housing of the gas turbine engine;wherein the airflow is drawn through the acoustic liner and exhaustedwithout use of a pump.
 9. The method of claim 8, and further comprisingdrawing the airflow from a base of the guide vane into the conduitdisposed in a core housing of the gas turbine engine.
 10. The method ofclaim 9, and further comprising: drawing the airflow into the pluralityof exhaust passages, wherein the plurality of exhaust passages extendfrom the conduit; and exhausting the airflow from the plurality ofexhaust passages to a cavity of the core housing to cool componentslocated in the cavity.
 11. The method of claim 9, and further comprisingdelivering the airflow to an active clearance control system in aturbine section of the gas turbine engine.
 12. The method of claim 11,and further comprising opening and closing a valve in the conduit toregulate cooling airflow to the turbine section for active clearancecontrol, wherein opening the valve allows the airflow to the turbinesection and closing the valve inhibits the airflow to the turbinesection and exhausts the airflow from a plurality of passages extendingfrom the conduit to a cavity of the core housing to cool componentslocated in the cavity.
 13. The method of claim 12, wherein the valve isclosed during a takeoff operation of the gas turbine engine and openduring a cruise operation of the gas turbine engine.
 14. A passivetranspirational flow acoustic liner assembly for use in a gas turbineengine, the passive transpirational flow acoustic liner assemblycomprising: a plurality of guides vane disposed in a fan section of thegas turbine engine, each guide vane comprising: an airfoil having asuction side and a pressure side; and a transpirational flow acousticliner disposed in the airfoil, the liner comprising: a face sheetdefining a portion of an outer surface of the airfoil and having aplurality of first apertures; a segmented member coupled to the facesheet, the segmented member comprising a plurality of chambers in fluidcommunication with the outer surface via the plurality of firstapertures; a backing sheet coupled to the segmented member such that thesegmented member is positioned between the face sheet and the backingsheet, the backing sheet having a plurality of second apertures; and aplenum coupled to the backing sheet opposite the segmented member, theplenum in fluid communication with the outer surface via the pluralityof second apertures, the plurality of chamber, and the plurality offirst apertures, wherein the plenum is open at a base of the airfoil andclosed at a tip of the airfoil; a manifold fluidly connected to theplenum of each guide vane by an opening at the base of the airfoil; anda conduit disposed in a core housing of the gas turbine engine andfluidly connected to the manifold, wherein the conduit comprises aplurality of exhaust passages extending from the conduit to a cavity inthe core housing, the plurality of exhaust passages located to deliverthe airflow to components disposed in the cavity.
 15. The assembly ofclaim 14, wherein the first apertures of the plurality of firstapertures have a diameter in a range of 3 to 50 thousandths of an inch.16. The assembly of claim 14, wherein the second apertures of theplurality of second apertures have a diameter in the range of 3 to 100thousandths of an inch.
 17. The assembly of claim 14, wherein the facesheet is located on the suction side of the airfoil.
 18. The assembly ofclaim 14, and further comprising: an active clearance control system ina turbine section of the gas turbine engine and fluidly connected to theconduit; and a valve disposed in the conduit in a fluid position aft ofthe plurality of exhaust passages and configured to regulate the airflowto the turbine section, wherein the airflow is delivered to the activeclearance control system when the valve is open and the airflow isdelivered to components disposed in the cavity of the core housing whenthe valve is closed.
 19. The assembly of claim 1, and furthercomprising: a plurality of guide vane assemblies; and a manifold fluidlyconnected to plenum of each guide vane assembly, wherein the conduit isfluidly connected to the manifold.
 20. The assembly of claim 3, whereinthe exhaust passages comprise a plurality of apertures directed towardthe components disposed in the core housing.